Turbine engines are used as the primary power source for various kinds of aircraft. Silicon-based substrates, such as silicon nitride and silicon carbide composites, are used for making turbine engine components that may be exposed to severe operating conditions. As used herein, “severe operating conditions” include high gas velocities (greater than about 500 m/second, exposure to salt, sulfur, and sand, and high temperatures of up to about 1500° C.). Although silicon-based substrates have good high temperature properties and many other advantages, they may be susceptible to recession under the severe operating conditions of an operating turbine engine. For example, water vapor may form inside the engine from gases in the atmosphere. The water vapor, when mixed with silicon from the component material, forms silicon hydroxide that can result in recession of silicon from the component, reducing component service life and necessitating its replacement or repair. Hot corrosion and erosion of the component from the salt, sulfur, and sand are also problems for substantially all substrates under severe operating conditions.
The resistance to recession of such silicon-based components, and the resistance to hot corrosion and erosion of silicon-based and non-silicon-based components, can be enhanced by applying protective coatings over the component. Conventional protective coatings include an environmental barrier coating (EBC), a thermal barrier coating (TBC), a braze layer, and combinations thereof. A “multi-layer protective coating” may be formed from the combination of coatings and/or layers. In a conventional multi-layer protective coating for a silicon-based component, the braze layer may be disposed between the silicon-based substrate or component and the environmental barrier coating (EBC), and the thermal barrier coating (TBC) may be disposed on the EBC. The braze layer substantially prevents diffusion of silicon nitride or silicon carbide from the silicon-based substrate or component into the EBC. Conventional braze layers contain silicon, tantalum, and chromium. The inclusion of tantalum and chromium reduces the melting point of the braze layer below that of a silicon braze layer and helps form additional intermetallic compounds that strengthen the braze layer and make it more resistant to oxidation and recession. However, such intermetallic compounds have different melting points making the coating operation difficult, especially for large scale components. For example, a chromium silicon braze has a melting point of 1305° C. and a tantalum silicon braze has a melting point of 1410° C.
The thermal barrier coating is typically a ceramic material such as zirconia or hafnia, and is stabilized with an oxide such as yttria to form yttria-stabilized zirconia or hafnia. The thermal barrier coating (TBC) effectively insulates the component from heat, reducing the temperature of the component and extending its service life. The TBC is itself susceptible to degradation by various processes that occur during operation of the turbine engine. One such degradation process that may occur is the formation of calcium-magnesium-aluminosilicate (CMAS) from engine dirt or other particles in the turbine engine. Under severe operating conditions, built-up CMAS on engine components may melt and penetrate pores in the TBC. As the built-up CMAS solidifies, the CMAS may cause stresses within the TBC, degrading the TBC and causing increased temperature and wear of the turbine engine components. Additionally, other chemical processes may occur as an indirect result of CMAS build-up, further degrading the TBC and damaging engine components.
In addition, mismatched coefficients of thermal expansions (CTE) between the substrate and overlying layer(s) of the protective coating induce stress in the coating. A silicon-based substrate typically has a low thermal expansion compared to overlying layer(s). To help compensate for the high CTE of the thermal barrier coating, for example, and reduce the CTE mismatch, porosity has been introduced into conventional TBCs by plastic spheres (typically polystyrene) that form pores when burned off at elevated temperatures. The increase in porosity decreases the CTE of the TBC to be closer to the CTE of the underlying layer(s). However, at severe operating conditions, at least a portion of the TBC may sinter, closing the pores. The reduction in porosity results in an increase in the CTE of the TBC, hereinafter referred to as a “sintering effect.” The sintering effect can result in stress-induced fractures of the coating.
Accordingly, it is desirable to provide a protective coating and a coated component comprising the protective coating. In addition, it is desirable to provide a protective coating with improved resistance to at least one of recession, corrosion, erosion, CMAS at severe operating conditions. It is also desirable to provide a protective coating that simplifies the coating operation, and reduces thermal expansion mismatches between the substrate and overlying layer(s) of the protective coating. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.